Gas turbine engines (“GTE”) have been engineered extensively to improve efficiency, thrust-to-weight ratios, and other measures of engine performance. One of the most direct manners by which engine performance can be improved is through increases in core rotational speeds and turbine inlet temperatures generated during engine operation. However, as turbine inlet temperatures and rotational speeds increase, so too do the thermal and mechanical demands placed on the GTE components. The most demanding performance requirements are typically placed on the high pressure turbine rotor or rotors, which are positioned immediately downstream of the GTE combustion section and which rotate at the greatest speed during engine operation. The turbine blades, in particular, are directly exposed to combustive gas flow at or near peak temperatures and are consequently heated to exceedingly high temperatures at which most alloys weaken or melt and become prone to oxidation or other forms of chemical degradation. By comparison, the inner annular body of the turbine (commonly referred to as the “turbine disk”) is largely shielded from direct exposure to combustive gas flow, but is subject to considerable mechanical stress resulting from both centrifugal forces acting on the turbine rotor at high rotational speeds and rim-to-bore thermal gradients.
Turbines can be broadly divided into two categories, axial and radial turbines, based upon the direction of airflow received by the turbine relative to the turbine's rotational axis. Compared to axial turbines, radial turbines offer certain performance benefits including superior pressure ratios; e.g., a single radial turbine is often capable of providing a power output equivalent to two similarly-sized axial turbines. Radial turbines also provide relatively smooth exit flow conditions as compared to axial turbine turbines, which are often characterized by highly turbulent outflow. However, by inherent design, radial turbines tend to be relatively lengthy in an axial direction and, thus, have lower power-to-weight ratios. In addition, it is difficult to fabricate radial turbines having internal cooling features, such as cooling flow passages formed within the turbine blades. Specifically, casting of radial wheels with internally-cooled blades has proven excessively costly due to low yields results from the usage of complex casting cores generally required to form the blade cooling passages. An uncooled radial turbine may be incapable of withstanding prolonged stresses at high operational speeds without premature fatigue in situations wherein a gas turbine engine is operated at relatively high combustive gas flow temperatures to boost power output. In such situations, the inability to manufacture an internally-cooled radial turbine in a reliable and cost-effective manner may prevent the usage of a radial turbine altogether and a number of cooled axial turbines may be employed instead. For at least these reasons, radial turbines tend to be utilized only within a relatively limited number of small, low-flow gas turbine engine platforms.
It is thus desirable to provide embodiments of a radial turbine suitable for usage in a gas turbine engine that is relatively lightweight and cost effective to implement. Ideally, embodiments of such a radial turbine would include internal cooling passages to permit operation of the radial turbine at higher temperatures. It would also be desirable if, in certain embodiments, the different sections of such a radial turbine were fabricated from disparate alloys tailored to the particular operating conditions experienced by each turbine section. Finally, it would be desirable to provide embodiments of a method for fabricating such a radial turbine. Other desirable features and characteristics of embodiments of the present invention will become apparent from the subsequent Detailed Description and the appended Claims, taken in conjunction with the accompanying drawings and the foregoing Background.